The present invention relates to an arrangement for an aerospace turbofan engine.
As the specific thrust of a conventional turbofan engine is reduced to improve propulsive efficiency, or to reduce fan and jet noise, its bypass ratio increases, the fan rotational speed reduces and the fan shaft torque increases, requiring a larger diameter shaft. The increased fan shaft diameter increases the minimum bore diameters of the core components, requiring them to be designed for lower rotational speeds that, in turn, increase parts counts, cost and weight. The increased bypass ratio also results in a mismatch between the hub radius of the fan and the hub radius of the subsequent compressor stages, increasing duct lengths and duct losses between compressors. The higher bypass ratio also requires an increased number of turbine stages to drive the fan, the maximum turbine radius being limited by installation constraints. Very high bypass ratio engines are prohibitively heavy, and when installed under-wing on a conventional low wing aircraft they require an increase in aircraft undercarriage length which adds further cost and weight penalties to the aircraft. The larger size engines are also more difficult to transport to and from the aircraft.
In order to avoid an excessive number of turbine stages and their associated cost, mass and complexity, the generally proposed solution for very high bypass ratio engines is to incorporate a gearbox between the fan and the turbine so that the turbine rotational speed can be increased to increase work per stage and reduce the total number of turbine stages. However this arrangement adds the cost, mass, complexity and potential unreliability of incorporating a high power gearbox, it is unproven for large engines, and is unattractive to aircraft operators. Problems still remain with the mismatch in compressor hub radii and with the installation of the engine on the aircraft.
Increasing the number of engines on an airframe by installing two or three smaller engines side by side in a common nacelle on each under-wing pylon is a well-known arrangement. This would enable lower specific thrust engines to be fitted without an increase in undercarriage length or a reduction in ground clearance or a reduction in the ‘gully depth’ (the distance between the wing and the nacelle) that is needed to minimise interference drag. Small engines are however less efficient and less cost effective than large engines and this arrangement does not solve the problem that each engine needs either a very large number of turbine stages or a geared fan.
Aircraft engine configurations where a single engine has more than one fan are also known.
Aircraft engines having two fans with parallel flows have been proposed with the fans arranged in series on the same shaft and driven by the same turbine. These so called tandem fan engine arrangements can increase the total mass flow for a given frontal area and reduce the total number of turbine stages required for a given overall bypass ratio. However, the convoluted exit nozzles for the first fan and the bifurcated intakes for the second fan add significantly to the installation losses for a low specific thrust version of such an engine. In a conventional under-wing installation the extra length of a tandem fan engine projecting forwards from under the wing results in higher wing bending moments and a significant weight penalty for the wing and the pylon supporting the engine. The forward extension of the engine also interferes with the space available for loading and unloading the aircraft. Variable specific thrust versions of such engines may be better suited to supersonic aircraft where the engine nacelles can be integrated with a delta wing.
The use of separate lift fans or propellers driven by shafts and gears through clutches is another known arrangement and multiple low specific thrust fans have also been proposed mounted around the rear fuselage of an aircraft and driven by gears and shafts from a smaller number of gas generators driving power turbines buried within the fuselage. These arrangements have the disadvantage of needing complex geared drive systems.
It has also been proposed to drive a low specific thrust fan by turbine blades mounted on the periphery of the fan rotor assembly. Such tip turbines can be fed from gas generators that are not coaxial with the fans. However, tip turbines are not easy to design and manufacture, and sealing them is particularly difficult. Gas leakages result in poor performance, making such arrangements unattractive.
Remote fans driven by compressor bleed air and separate combustors and turbines have also been proposed. Such engines need additional combustors that add to the complexity and potential unreliability of the engines, making them unattractive. The fans would also need to be geared to avoid an excessive number of turbine stages in a low specific thrust power plant.
A turbojet engine with an auxiliary pair of parallel flow fans arranged on either side of it and driven by parallel flow turbines is also known. Such a multiple fan aero engine arrangement was first described in GB1,110,113. This engine has a turbojet core with its own intake, and two auxiliary parallel flow fans with their own separate intakes. An auxiliary turbine directly drives each of the auxiliary fans. The core exhaust gasses can be supplied to the two parallel flow auxiliary turbines by means of a bifurcated duct. This engine, mounted with the fans side by side in an under-wing installation, overcomes the problem with ground clearance for low specific thrust engines. The disadvantages of this multiple fan engine arrangement, for a low specific thrust engine, are that a very large number of turbine stages are required; the core intake is not protected from ingestion of foreign objects by an upstream fan; the core needs its own intake with its additional pressure losses; and the core compressors do not benefit from the supercharging of the air through a fan.
None of these previously described designs is entirely satisfactory for a very low specific thrust engine mounted in an under-wing installation.